(1) Field of the Invention
The present invention relates to the field of the emergency piloting of aircraft.
(2) Description of Related Art
Emergency piloting means the pilot maintaining a certain amount of authority over changes in the attitude of the aircraft in the event of one or more of the manual flight control systems that are provided for controlling such changes of attitude in normal operation not operating normally.
Without this being a limitation, the invention is described in its applications to rotary wing aircraft, such as helicopters and other rotorcraft or hybrid craft. The invention also applies to other aircraft having so-called “fixed” wings (airplanes, gliders, or the like).
The flight control systems concerned by the invention are “manual”. Consequently, in order to obtain a change in the flight attitude of the aircraft in normal operation, the pilot acts on a control device (e.g. a stick, a lever, or a pedal—cf. below) that forms part of one or more manual flight control systems, and acts thereon.
These manual flight control systems include one or more airfoil surfaces such as the blades of one or more rotors and/or other surfaces (wings, flaps, fins, stabilizers, rudders, etc.) and they are designed to control the actions thereof. It is these airfoil surfaces themselves that actually generate the changes in attitude of the aircraft.
Specifically, a manual flight control system transmits authority from the pilot for changing the flight attitude of the aircraft in roll, in pitching, or in yaw.
Naturally, the term “manual” should not be taken in its literal meaning. It is used specifically to distinguish between a flight control system operated by a human and an automatic control system (e.g. an autopilot or automatic flight control system (AFCS)). Specifically, a control device such as the mechanical system for controlling yaw, which is usually actuated by the feet acting on pedals, should also be considered as a “manual” flight control system.
Although a flight control system is said to be “manual”, it may also include power assistance, in order to help the pilot move the manual control device against external reaction forces from the airfoil surfaces.
It is not uncommon for a manual flight control system also to restore a certain amount of “feel” to the pilot representative of these reaction forces, by applying forces to the manual control devices. The term “haptic” or “TRIM” is used to describe such feedback that acts relative to the trim of the control system.
For transmitting movements from upstream to downstream, a manual flight control system thus comprises at least one manual flight control device (at the upstream end), intermediate mechanical connections referred to as “linkages” for transmitting control, and at the downstream end an airfoil surface such as a rotary wing (e.g. rotor blades).
In order to illustrate the description of manual flight control systems, reference is made to a conventional helicopter.
In general, a manual flight control system known as the “cyclic pitch” system acts on the cyclic pitch of the blades of the main rotor so as to cause a change in flight attitude in roll and in pitching, by actuating a corresponding manual flight control device known as the “cyclic stick”.
Another manual flight control system known as the “collective pitch” system acts on the collective pitch of the blades in order to give rise to a vertical change in flight attitude (i.e. a change in altitude) of the aircraft, by acting on a corresponding manual flight control device known as the “collective pitch” lever or the “collective” lever.
Yet another manual flight control system is dedicated to changing flight attitude in yaw, and it is actuated by the pilot using a flight control device in the form of pedals.
It should be observed that a mixer enables collective pitch controls and cyclic pitch controls to operate independently of each other and without mutual interaction.
By way of example, in a heavy helicopter, a manual flight control system often includes a phasing unit; and in a helicopter having a “tail” anti-torque rotor, the yaw manual flight control system comprises not only pedals (control device) and primary and secondary (yaw) linkages, but also means for coupling collective pitch and yaw, and also a mixer.
These manual flight control systems are thus relatively long and heavy, in particular for transmitting movements to control changes of attitude in yaw.
Such systems generate friction forces that can be considerable, particularly when the systems are long and/or complex (i.e. include numerous components).
In theory, these forces could lead to electrical flight controls being used, as described in documents US 2004/200928 or U.S. Pat. No. 7,229,046. Nevertheless, such electrical flight controls are difficult to implement in practice. In particular on existing aircraft they give rise to modifications that are extensive and expensive.
Faced with this difficulty, a practical solution is to provide hydraulic or pneumatic power assistance in the manual flight control system.
Other documents are mentioned below that warrant being considered in association with the invention.
In order to provide versatile piloting assistance, document FR 2 946 620 describes a movement transmission system in which an additional force is generated using piloting assistance means that are mechanically associated with said system, and that operate as a function of the position of the flight control and of an instantaneous force as measured on the movement transmission system. Autopilot means can control the piloting assistance in order to make it perform the functions of existing dampers and parallel actuators so as to enable it to take the place of the parallel actuators, the dampers, and the hydraulic unit that would conventionally be used. The piloting assistance may optionally include at least one series actuator arranged between the rotor and the connection means.
Document FR 2 931 132 describes assisted flight controls for a rotorcraft. A control member such as a stick is connected to crank means including an incorporated motor having a drive rotor and a stator, the motor assisting flight control movements as a function of measurements from piloting movement sensors.
Document FR 2 407 130 describes redundant flight control for a helicopter, having flexible cables making it possible to selectively position a desired flight quadrant. Short balancing springs are used for putting the quadrant in a preselected position in the event of the cables breaking.
Document FR 2 912 375 describes an example of a “series” actuator, specifically a smart electromechanical actuator (SEMA). That actuator comprises a rotary electric motor, position sensors, and a servo-control circuit connected to the two position sensors and to the motor in order to deliver a motor power signal that varies as a function of a position setpoint signal and as a function of the signals delivered by the two position sensors. Redundant modules calculate a position as a function of the two sensors, under mutual monitoring of the results delivered by those calculation modules, in order to produce a fault detection signal.
Document FR 2 920 744 describes an electric compensation actuator for actuating the controls of a rotorcraft. An autopilot system transmits control signals to electric actuators connected in series within the flight controls. A compensation actuator is also fitted to the controls, in particular the pitching and roll controls, in order to maintain an anchor point (or “middle” point) for the manual control device (in particular the “cyclic stick”). The compensation actuator comprises an (auto)synchronous rotary motor and reversible speed-reducing gearing. A motor control circuit varies the gear ratio of reduction stages, as a function of an applied position setpoint.
Document FR 2 946 317 describes a crank for a manual flight control device of a rotary wing aircraft that serves to apply variable gain to a system for transmitting movement and in order to change flight attitude.
Document U.S. Pat. No. 4,492,907 describes a yaw control system for a helicopter, with a connection system having hydraulic power assistance and two servo-controls acting in parallel with the connection system while in automatic flight mode. A position servo-control circuit includes a circuit for controlling the position of the member that is driven by a series servo-motor, and that may in particular be measured by means of a potentiometer.
Document U.S. Pat. No. 4,529,155 describes a flight control system for a helicopter with semicircular quadrants and floating end fittings each having a flexible cable anchored thereon. Those floating end fittings are biased by springs to accommodate the event of a cable separating. In the event of the opposite cable breaking, the springs perform a centering action.
Document US 2010/072322 describes a pedal system for yaw piloting with little movement. In order to act on an anti-torque rotor, there are provided the pedal system, a servo-control, and interposed control logic.
Although they are of interest, those documents are in practice poorly adapted to the very particular context of emergency piloting in the event of a break in a flight control of a rotary wing aircraft, e.g. a helicopter.
In this context, the invention seeks to solve the following technical problems, in particular.
For obvious reasons of safety in aviation, various solutions have been proposed to mitigate a breakage of the flight controls of an aircraft. Such a breakage within a manual flight control system could lead to the aircraft being lost.
For example, a manual flight control system for controlling a helicopter in yaw is known that is fitted with a spring-loaded centering rod, as is to be found in particular on the following Eurocopter helicopters: EC330, EC332, and EC225.
Such a spring and centering rod system presents functions that are comparable to the subject matter of document U.S. Pat. No. 4,529,155. By imposing mechanical equilibrium, the system makes it possible to “freeze” the position of the airfoil surface that acts like a rudder (i.e. the anti-torque rotor in the yaw system), in a position that is predetermined for use in the event of the yaw control breaking.
Although that spring and centering rod system is effective, its return position is constant and therefore corresponds to a single flight setting only, and, in accordance with current procedures, it must be associated with a “refuge” speed. In other words, the pilot can no longer impose a yaw change of attitude on the aircraft, and yet the pilot must ensure that the aircraft is at an imposed speed on its trajectory (referred to as the “refuge” speed, i.e. a speed that is standardized for a given aircraft, e.g. equal to a maximum power speed).
By way of example, this can make it necessary to perform a landing with forward speed: the aircraft touches the ground while advancing at a certain speed. For a rotary wing aircraft that does not have a wheeled undercarriage, that will give rise, at best, to a “hard” landing.
Other approaches, such as that used on “Black Hawk” H60 aircraft, are substantially similar to that described above. Nevertheless, those approaches tend to increase the control forces that need to be applied to the manual flight control system, in particular while it is being controlled in normal mode. These forces need to be overcome both by a series actuator (similar to the subject matter of document FR 2 912 375) and by the parallel or “trim” actuator, so there is no possibility of piloting in the event of a linkage (cable) breaking, with only one equilibrium point between control force and control stroke, and with a given value for the return force.